Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts 26, 28, 30.
In view of the above it will be appreciated that the performance of a gas turbine engine cycle, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. In such circumstances it is desirable to operate the gas turbine at the highest possible gas temperature. For any engine cycle, compression ratio or bypass ratio, increasing the turbine entry gas temperature will always produce more specific thrust but as turbine entry temperatures increase it will also be understood that the life of an uncooled turbine blade falls. In order to meet these increased turbine entry temperatures it is therefore necessary to develop better materials and to introduce internal cooling air.
In modern gas turbine engines the high pressure turbine gas temperature is generally now much hotter than the melting point of the material used and in some engine designs the intermediate pressure and low pressure turbines are also cooled to remain within acceptable operational parameters particularly for life expectancy. During passage through the gas turbine engine the mean temperature of the gas flow stream decreases as power is extracted so the need to cool static and rotating parts of the engine decreases as the gas moves from the high pressure stages through the intermediate and low pressure stages towards an exit nozzle.
It is known to utilise internal convection and external coolant films as methods for cooling in gas turbine engines. In such circumstances high pressure turbines and nozzle guide vanes (NGVs) consume relatively large amounts of cooling air on high temperature parts of engines. High pressure blades typically use about half the cooling air that is required for the nozzle guide vanes. The intermediate and low pressure stages downstream of the high pressure turbine use progressively less cooling air.
FIG. 2 provides an isometric view of a typical single stage cooled high pressure turbine arrangement including a nozzle guide vane assembly 31 and a high pressure turbine blade assembly 30. The nozzle guide vane assembly 31 includes guide vanes 32 presented between an inner platform 33 and an outer platform 34. The high pressure turbine rotor blade assembly comprises blades 35 extending from platforms 36 secured through roots 37 to a rotor assembly 38. At an outer end 39 of the blades 35 shrouds are provided to limit gas flow leakage.
Cooling of the blades 35 and the guide vanes 32 is achieved through use of high pressure air bleed from a compressor (not shown). Part of the high pressure air flow from the compressor bypasses the combuster and is therefore relatively cool compared to the gas temperature driving the blades 35 and guided by the aerofoil 32. Typically the temperatures will be in the order of 700 to 1,000K whilst the gas temperatures presented to the vanes 32 and the blades 35 will be in excess of 2,100K. It will be understood that the cooling air from the compressor (not shown) utilised to cool the hot turbine components is not utilised to produce work from the turbine and so the engine. In such circumstances the coolant flow represents lost power and therefore has an adverse effect upon overall engine operating efficiency. Thus it is important to utilise the cooling air as effectively as possible.
Previously it is known to provide cooling effects with respect to the high pressure turbine rotor blades using a combination of internal convective cooling and external film cooling. The leading edge portion of turbine blades is therefore cooled by such processes and utilises either augmented channel flow or impingement convective cooling plus film cooling in the region of the stagnation point for the blade.
Impingement cooling is considered superior to augmented channel flow and is favoured when dealing with modern engine applications running at elevated gas temperatures as illustrated above. However, the important peak heat transfer coefficient levels associated with impingement jet cooling are only attainable when adequate pressure ratios are achieved across the jets. The pressure ratios drive the required cooling flow levels through the jets to keep the Reynolds numbers as high as physically possible within overall design constraints. It will be appreciated the design constraints that limit the impingement jet cooling performance include coolant feed pressure upstream of the jets and local gas path static pressure distribution on the external surface of the respective aerofoils defining the turbine blades such as in the vicinity of the aerofoil leading edge.
FIG. 3 provides a schematic cross sectional view through an aerofoil blade 41 with an impingement cooled leading edge arrangement 42. It will be appreciated that coolant flows pass in the direction of the arrowheads depicted. In such circumstances the coolant passes radially up an augmented feed passage 43 towards a tip of the blade 41. A series of impingement jets progressively bleed coolant through apertures 44 (i.e., 44a, 44b and 44c) across a divider wall into a number of individual impingement plenum chambers or cavities 45 (i.e., 45a, 45b and 45c) aligned radially up the leading edge of the blade 41. These cavities 45 are typically referred to as Boxcars and act as pleniums from which the leading edge film cooling flow is bled under pressure out of outlet apertures (i.e., 46a, 46b and 46c) to provide a film cooling effect 47 on an external surface of the aerofoil defining the blade. The pressure in the chambers or cavities 45 is kept at a level suitably above that of the local gas flow about the blades in order to ensure that hot gas ingestion never occurs under adverse operating conditions even when the engine and in particular the blades are nearing the end of their useful life.
In the above circumstances the impingement jet pressure ratio is virtually fixed along with the quantity of coolant that can be presented across the apertures 44, 46 for a given design of aerofoil in a blade 41. The level of transferred coolant air through the jets or apertures 44, 46 is therefore also virtually fixed unless pressure can be increased to the blade. Unfortunately, increasing the pressure to the blade can only be achieved at the expense of engine performance and is limited due to increased leakage (FIG. 2) and work extraction pumping the air up the front face of the disc to the blade feed passage 43. As can be seen in FIG. 3 the apertures 44 are generally angled with respect to the feed passage flow in such a manner that a proportion of the dynamic pressure head in addition to the static pressure is utilised to drive an impingement flow A across the cavities 45 to the apertures 46. Such an approach helps maximise available pressure ratio across to the apertures or jets 44. It will be appreciated that the inflows to cavity 45 through the apertures 44 must equal the outflow from that cavity through the outlet apertures 46.
In FIG. 3 the pressure in the feed passage 43 will generally increase as it flows up the blade from root to tip due to a centrifugal effect of rotation. In such circumstances, rotation provides a pumping effect which results in the feed pressure being higher at the entrance to the apertures 44 at the outer parts 44c compared to the feed pressure for the inner cavities 45 through inner apertures 44a and 44b. Furthermore, the external static pressure distribution also rises from the root to the tip of the blade but not as much as that internal pressure and consequently the pressure ratio across the apertures 44 rises from the root to the tip up the leading edge of the blade 41. Such increases in the pressure ratio will lead to levels of impingement heat transfer which also rise further up the leading edge of the blade 41. However, the heat load experienced by the blade 41 leading edge generally peaks at approximately mid span due to the radial gas temperature distribution originating from the combustor. This heat distribution is difficult to accommodate with previous cooling arrangements.
It will also be appreciated that in addition to the effects described above radial stress distribution will tend to be higher at the root sections and lower at the tip sections of the blade 41 due to the centrifugal loading on the aerofoil of the blade 41. Therefore, there is typically a need to cool the lower and mid portion of the aerofoil of the blade 41 more than the tip to retain structural integrity. However, as indicated, the internal cooling due to the pressure differential as a result of rotation is generally more effective at the tip of the blade 41.